Turbine bucket assembly and methods for assembling same

ABSTRACT

A method for assembling a rotor assembly for use with a turbine engine. The method includes providing at least two rotor blades that each include a shank extending between a dovetail and a platform. The shank includes at least one cover plate that extends inwardly from the platform towards the dovetail. An airfoil extends outwardly from the platform. A first rotor blade is coupled to a rotor disk. A second rotor blade is coupled to the rotor disk, such that a cavity is defined between the first and second rotor blades, and such that a seal path is defined between a first rotor blade cover plate and a second rotor blade cover plate.

BACKGROUND OF THE INVENTION

The subject matter described herein relates generally to gas turbineengines and, more particularly, to a bucket assembly for use with aturbine engine.

At least some known rotor assemblies used with turbine engines includeat least one row of circumferentially-spaced rotor blades. Each rotorblade includes an airfoil that includes a pressure side and a suctionside that are connected together along leading and trailing edges. Eachairfoil extends radially outward from a rotor blade platform. Each rotorblade also includes a dovetail that extends radially inward from a shankdefined between the platform and the dovetail. The dovetail is used tomount the rotor blade to a rotor disk or spool. Known blades are hollowand include an internal cooling cavity that is defined at leastpartially by the airfoil, platform, shank, and dovetail and that is usedto channel a flow of cooling fluid. Leakage of cooling fluid may occurbetween adjacent rotor blades. Depending on the amount of leakage,turbine performance and output may be adversely impacted.

Furthermore, the airfoil portions of at least some known rotor bladesare generally exposed to higher temperatures than the dovetail portions.Higher temperatures may cause temperature mismatches to develop at theinterface between the airfoil and the platform, and/or between the shankand the platform. These temperature mismatches may cause compressivethermal stresses to be induced to the rotor blade platform. Over time,continued operation with high compressive thermal stresses may causeplatform oxidation, platform cracking, and/or platform creep deflection,any or all of which may shorten the useful life of the rotor assembly.

BRIEF SUMMARY OF THE INVENTION

In one aspect, a method for assembling a rotor assembly for use with aturbine engine is provided. The method includes providing at least tworotor blades that each include a shank extending between a dovetail anda platform. The shank includes at least one cover plate that extendsinwardly from the platform towards the dovetail. An airfoil extendsoutwardly from the platform. A first rotor blade is coupled to a rotordisk. A second rotor blade is coupled to the rotor disk, such that acavity is defined between the first and second rotor blades, and suchthat a seal path is defined between a first rotor blade cover plate anda second rotor blade cover plate.

In a further aspect, a rotor blade for a turbine engine is provided. Therotor blade includes a platform that includes a radially outer surfaceand a radially inner surface. An airfoil extends radially outwardly fromthe platform. A dovetail is adapted to be coupled to a rotor wheel. Ashank extends between the platform and the dovetail. The shank includesat least one cover plate that extends inwardly from the platform towardsthe dovetail. At least one sealing assembly is coupled to the coverplate. The sealing assembly extends from the dovetail to the platform.The sealing assembly forms a seal path between the rotor blade and acircumferentially adjacent rotor blade.

In another aspect, a gas turbine engine is provided. The gas turbineengine includes a compressor and a combustor coupled downstream from thecompressor to receive at least some of the air discharged by thecompressor. A rotor shaft is coupled to the compressor. A plurality ofcircumferentially-spaced rotor blades are coupled to the rotor shaft.Each of the plurality of rotor blades includes a platform. An airfoilextends radially outwardly from the platform. A dovetail is coupled tothe rotor shaft. A shank extends between the platform and the dovetail.The shank includes at least one cover plate that extends inwardly fromthe platform towards the dovetail. At least one sealing assembly iscoupled to the cover plate such that a seal path is defined betweenadjacent rotor blades.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is schematic illustration of an exemplary known turbine enginesystem.

FIG. 2 is an enlarged perspective view of an exemplary rotor assemblythat may be used with the turbine engine system shown in FIG. 1.

FIG. 3 is an enlarged sectional view of a portion of the rotor assemblyshown in FIG. 2

FIG. 4 is a cross-sectional view of the rotor assembly shown in FIG. 2.

DETAILED DESCRIPTION OF THE INVENTION

The exemplary methods and systems described herein overcomedisadvantages of known rotor blade assemblies by providing a rotor bladethat facilitates reducing leakage of cooling fluid from the rotor blade.More specifically, the embodiments described herein include a labyrinthseal path that is positioned between adjoining rotor blades tofacilitate increasing a back pressure between adjacent rotor blades andto facilitate reducing leakage of cooling fluid through the rotorblades.

As used herein, the term “rotor blade” is used interchangeably with theterm “bucket” and thus can include any combination of a bucket includinga platform and dovetail and/or a bucket integrally formed with the rotordisk, either of which may include at least one airfoil segment.

FIG. 1 is a schematic view of an exemplary gas turbine engine 10. In theexemplary embodiment, gas turbine engine 10 includes an intake section12, a compressor section 14 coupled downstream from intake section 12, acombustor section 16 coupled downstream from compressor section 14, aturbine section 18 coupled downstream from combustor section 16, and anexhaust section 20. Turbine section 18 is includes a rotor assembly 22that is coupled to compressor section 14 via a drive shaft 32. Combustorsection 16 includes a plurality of combustors 24. Combustor section 16is coupled to compressor section 14 such that each combustor 24 is inflow communication with compressor section 14 and such that fuel nozzleassembly 26 is coupled to each combustor 24. Turbine section 18 isrotatably coupled to compressor section 14 and to a load 28 such as, butnot limited to, an electrical generator and a mechanical driveapplication. In the exemplary embodiment, compressor section 14 andturbine section 18 each include at least one turbine blade or bucket 30coupled to rotor assembly 22 that include airfoil portions (not shown inFIG. 1).

During operation, intake section 12 channels air towards compressorsection 14. Compressor section 14 compresses the inlet air to a higherpressure and temperature and discharges the compressed air towardscombustor section 16. The compressed air is mixed with fuel and ignitedto generate combustion gases that flow to turbine section 18. Turbinesection 18 drives compressor section 14 and/or load 28. Specifically, atleast a portion of compressed air supplied to fuel nozzle assembly 26.Fuel is channeled to fuel nozzle assembly 26 wherein it is mixed withthe air and ignited in combustor section 16. Combustion gases aregenerated and channeled to turbine section 18 wherein gas stream thermalenergy is converted to mechanical rotational energy. Exhaust gases exitturbine section 18 and flow through exhaust section 20 to ambientatmosphere.

FIG. 2 is an enlarged perspective view of an exemplary rotor assembly 22that may be used with gas turbine engine 10 (shown in FIG. 1). FIG. 3 isan enlarged sectional view of a portion of rotor assembly 22, and FIG. 4is a cross-sectional view of rotor assembly 22 taken along sectionalline 4-4 in FIG. 3. In the exemplary embodiment, rotor assembly 22includes at least one rotor blade 100 coupled to a rotor disk 102.Moreover, in the exemplary embodiment, rotor assembly 22 includes afirst rotor blade 104, a second rotor blade 106, and at least a thirdrotor blade 107. In the exemplary embodiment, each rotor blade 100 iscoupled to a rotor disk 102 that is rotatably coupled to a rotor shaft,such as drive shaft 32 (shown in FIG. 1). In an alternative embodiment,rotor blades 100 are mounted within a rotor spool (not shown). Morespecifically, when rotor blades 100 are coupled to rotor disk 102, a gap108 is defined between adjacent circumferentially-spaced rotor blades100. In the exemplary embodiment, each rotor blade 100 extends radiallyoutward from rotor disk 102 and includes an airfoil 110, a platform 112,a shank 114, and a dovetail 116. Each airfoil 110 includes a firstsidewall 118 and a second sidewall 120 that is coupled to first sidewall118 to form airfoil 110.

In the exemplary embodiment, first sidewall 118 is convex and defines asuction side 119 of airfoil 110, and second sidewall 120 is concave anddefines a pressure side 121 of airfoil 110. First sidewall 118 iscoupled to second sidewall 120 along a leading edge 122 and along anaxially-spaced trailing edge 124 of airfoil 110. More specifically,airfoil trailing edge 124 is spaced chord-wise and downstream fromairfoil leading edge 122. First sidewall 118 and second sidewall 120each extend longitudinally or radially outwardly in span from a bladeroot 126 positioned adjacent to platform 112, to an airfoil tip 128. Inthe exemplary embodiment, an internal cooling chamber 130 is definedwithin airfoil 110 between first sidewall 118 and second sidewall 120,and extends through platform 112, through shank 114, and into dovetail116.

Platform 112 extends between airfoil 110 and shank 114 such that eachairfoil 110 extends radially outwardly from platform 112. Shank 114extends radially inwardly from platform 112 to dovetail 116. Dovetail116 extends radially inwardly from shank 114 to enable rotor blades 100to be coupled to rotor disk 102. Platform 112 includes an upstream sideor skirt 132, and a downstream side or skirt 134 that are connectedtogether with a pressure-side edge 136 and an opposite suction-side edge138. When rotor blades 100 are coupled to rotor disk 102, a gap 140 isdefined between circumferentially adjacent rotor blade platforms 112,and more specifically between pressure-side edge 136 and an adjacentsuction-side edge 138.

In the exemplary embodiment, shank 114 includes a first sidewall 142, asecond sidewall 144, an upstream sidewall or forward cover plate 146,and an opposite downstream sidewall or aft cover plate 148. Moreover, inthe exemplary embodiment, first sidewall 142 is substantially concaveand is coupled between forward cover plate 146 and aft cover plate 148such that forward cover plate 146 is opposite aft cover plate 148.Second sidewall 144 is substantially convex and is coupled betweenforward cover plate 146 and aft cover plate 148. In one embodiment,first sidewall 142 is coupled to second sidewall 144 such that a cavity150 is defined at least partially between first sidewall 142 and secondsidewall 144. In an alternative embodiment, first sidewall 142 iscoupled to second sidewall 144 such that a unitary member extendingbetween forward cover plate 146 and aft cover plate 148 is formed. Inanother alternative embodiment, shank 114 is formed as a unitary member.In the exemplary embodiment, first sidewall 142 and second sidewall 144are each recessed with respect to forward cover plate 146 and aft coverplate 148, respectively, such that when rotor blades 100 are coupled torotor disk 102, a shank cavity 152 is defined between first sidewall 142and an adjacent second sidewall 144.

In the exemplary embodiment, a forward angel wing 154 extends outwardlyfrom forward cover plate 146. An aft angel wing 156 extends outwardlyfrom aft cover plate 148. Forward angel wing 154 and aft angel wing 156each facilitate sealing forward and aft angel wing buffer cavities (notshown) defined within rotor assembly 22. In addition, a forward lowerangel wing 158 extends outwardly from forward cover plate 146, and isconfigured to facilitate sealing between rotor blade 100 and rotor disk102. More specifically, forward lower angel wing 158 extends outwardlyfrom forward cover plate 146 between dovetail 116 and forward angel wing154.

In the exemplary embodiment, aft cover plate 148 includes a leading edgeportion 164 and a circumferentially-spaced trailing edge portion 166. Afirst sealing assembly 168 is coupled to leading edge portion 164, and asecond sealing assembly 170 is coupled to trailing edge portion 166. Inthe exemplary embodiment, first sealing assembly 168 cooperates with anadjacent second sealing assembly 170 when rotor blades 100 are coupledto rotor disk 102. First sealing assembly 168 and second sealingassembly 170 each extend between dovetail 116 and platform 112, and eachfacilitates sealing shank cavity 152. In the exemplary embodiment, firstsealing assembly 168 and second sealing assembly 170 cooperate to form aseal path 172 between a first aft cover plate 148 and an adjacent secondaft cover plate 148. Seal path 172 facilitates reducing a volume of airchanneled between circumferentially adjacent rotor blade shanks 114.More specifically, seal path 172 facilitates reducing the volume of airthat must be channeled from forward cover plate 146 to aft cover plate148 through shank cavity 152 to facilitate preventing a flow of hotgases from entering shank cavity 152.

In the exemplary embodiment, aft cover plate 148 extends a radial heightr₁ from dovetail 116 to a platform inner surface 174. First sealingassembly 168 and second sealing assembly 170 each extend a radial heightr₂ from dovetail 116 to platform inner surface 174. Radial height r₂ isapproximately the same height as radial height r₁ of aft cover plate148. In one embodiment, first sealing assembly 168 and/or second sealingassembly 170 extends the full radial height r₁ of aft cover plate 148.

In one embodiment, first sealing assembly 168 includes a sealingextension 176 that extends outwardly from leading edge portion 164towards an adjacent rotor blade trailing edge portion 166. Secondsealing assembly 170 includes a recessed sealing groove 178 that isdefined within trailing edge portion 166. Recessed sealing groove 178 issized to receive an adjacent sealing extension 176 such that recessedsealing groove 178 and sealing extension 176 cooperate to form seal path172. In an alternative embodiment, first sealing assembly 168 includesrecessed sealing groove 178 and second sealing assembly 170 includessealing extension 176.

In the exemplary embodiment, first rotor blade 104 includes firstsealing assembly 168, including sealing extension 176, and secondsealing assembly 170, including recessed sealing groove 178. In analternative embodiment, first rotor blade 104 includes first sealingassembly 168, including recessed sealing groove 178, and second sealingassembly 170, including a sealing extension 176. In one embodiment,second rotor blade 106 includes first sealing assembly 168 and secondsealing assembly 170 each including sealing extension 176. In analternative embodiment, second rotor blade 106 includes first sealingassembly 168 and second sealing assembly 170 each including recessedsealing groove 178.

In the exemplary embodiment, recessed sealing groove 178 includes aradially outer surface 184 that extends between dovetail 116 andplatform inner surface 174. An abradable layer 186 is coupled torecessed sealing groove outer surface 184. Alternatively, in oneembodiment, abradable layer 186 includes an aluminum composite material.In the exemplary embodiment, sealing extension 176 includes a pluralityof labyrinth teeth 188 that extend outwardly from an inner surface 190of sealing extension 176. Labyrinth teeth 188 are each positionedadjacent to an opposing recessed sealing groove outer surface 184 suchthat a labyrinth seal 191 is defined between sealing extension 176 andrecessed sealing groove 178.

In the exemplary embodiment, shank 114 includes a leading edge radialseal pin slot 192 that extends generally radially through shank 114 atleast partially between platform 112 and dovetail 116. Morespecifically, leading edge radial seal pin slot 192 is defined withinshank forward cover plate 146 and is adjacent to shank convex sidewall144. Leading edge radial seal pin slot 192 is sized to receive a radialseal pin 194 to facilitate sealing between adjacent forward cover plates146 when rotor blades 100 are coupled within rotor disk 102. In oneembodiment, radial seal pin 194 is not inserted into leading edge radialseal pin slot 192. In an alternative embodiment, forward cover plate 146includes a first sealing assembly 168 and a second sealing assembly 170.

Referring to FIG. 3, in the exemplary embodiment, during operation ofgas turbine engine assembly 10, combustor section 16 generates andchannels combustion gases, represented by arrows 196, to rotor assembly22. Combustion gases 196 contact rotor blades 100 causing rotor assembly22 to rotate about drive shaft 32. At least a portion of combustiongases 196 pass through adjacent forward cover plates 146, around radialseal pin 194, and into shank cavity 152. First sealing assembly 168 andsecond sealing assembly 170 each facilitate preventing combustion gases196 from passing through adjacent aft cover plates 148 causing anincrease in a fluid pressure within shank cavity 152 that facilitatesreducing a volume of combustion gases 196 entering shank cavity 152.

The above-described methods and apparatus facilitate reducing anoperating temperature of a rotor assembly. More specifically, thelabyrinth seal defined between adjacent rotor blades facilitate reducingleakage of cooling fluid between adjacent rotor blades. In addition, theembodiments described herein facilitate increasing a back pressure ofcooling fluid within a shank cavity, which facilitates increasing a flowof cooling fluid to the rotor blades to reduce an operating temperatureof the rotor assembly. As such, the cost of maintaining the gas turbineengine system is facilitated to be reduced.

Exemplary embodiments of methods and apparatus for a turbine bucketassembly are described above in detail. The methods and apparatus arenot limited to the specific embodiments described herein, but rather,components of systems and/or steps of the method may be utilizedindependently and separately from other components and/or stepsdescribed herein. For example, the methods and apparatus may also beused in combination with other combustion systems and methods, and arenot limited to practice with only the gas turbine engine assembly asdescribed herein. Rather, the exemplary embodiment can be implementedand utilized in connection with many other combustion systemapplications.

Although specific features of various embodiments of the invention maybe shown in some drawings and not in others, this is for convenienceonly. Moreover, references to “one embodiment” in the above descriptionare not intended to be interpreted as excluding the existence ofadditional embodiments that also incorporate the recited features. Inaccordance with the principles of the invention, any feature of adrawing may be referenced and/or claimed in combination with any featureof any other drawing.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they have structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. A method for assembling a rotor assembly for usewith a turbine engine, said method comprising: providing at least tworotor blades that each include a shank extending between a dovetail anda platform, wherein each shank includes at least one cover plate thatextends inwardly from the platform towards the dovetail, and an airfoilthat extends outwardly from the platform; coupling a first rotor bladeto a rotor disk; coupling a second rotor blade to the rotor disk, suchthat a cavity is defined between the first and second rotor blades, andsuch that a seal path is defined between a first rotor blade cover plateand a second rotor blade cover plate.
 2. A method in accordance withclaim 1, further comprising: coupling a first sealing assembly to thefirst rotor blade cover plate; and coupling a second sealing assembly tothe second rotor blade cover plate to form a labyrinth seal path.
 3. Amethod in accordance with claim 2, further comprising: coupling asealing extension to the first rotor blade cover plate to form the firstsealing assembly; and defining a sealing groove within the second rotorblade cover plate.
 4. A method in accordance with claim 3, furthercomprising coupling an abradable surface to an outer surface of thesealing groove.
 5. A method in accordance with claim 3, furthercomprising coupling a plurality of labyrinth teeth to the sealingextension such that a tortuous path is defined between the first sealingassembly and the second sealing assembly.
 6. A method in accordance withclaim 2, wherein the first sealing assembly extends between the platformand the dovetail.
 7. A rotor blade for a turbine engine, said rotorblade comprising: a platform comprising a radially outer surface and aradially inner surface; an airfoil extending radially outwardly fromsaid platform; a dovetail adapted to be coupled to a rotor wheel; ashank extending between said platform and said dovetail, said shankcomprising at least one cover plate extending inwardly from saidplatform towards said dovetail; and at least one sealing assemblycoupled to said cover plate, said sealing assembly extending from saiddovetail to said platform, said sealing assembly forms a seal pathbetween said rotor blade and a circumferentially adjacent rotor blade.8. A rotor blade in accordance with claim 7, wherein said sealingassembly comprises a sealing extension coupled to said cover plate, saidsealing extension extending outwardly from said cover plate towards anadjacent rotor blade.
 9. A rotor blade in accordance with claim 8,wherein said sealing extension comprises a plurality of labyrinth teethextending outwardly from said sealing extension, said labyrinth teethconfigured to form a tortuous path between said sealing extension and anadjacent rotor blade.
 10. A rotor blade in accordance with claim 7,wherein said sealing assembly comprises a recessed sealing groovedefined within said cover plate.
 11. A rotor blade in accordance withclaim 10, wherein said sealing groove comprises an abradable surfaceextending from an outer surface of said sealing groove.
 12. A rotorblade in accordance with claim 7, further comprising a first sealingassembly coupled to said cover plate and an opposite second sealingassembly coupled said cover plate.
 13. A rotor blade in accordance withclaim 12, wherein said first sealing assembly comprises a sealingextension, said second sealing assembly comprises a recessed groove. 14.A rotor blade in accordance with claim 12, wherein said first sealingassembly and said second sealing assembly each comprise a sealingextension.
 15. A rotor blade in accordance with claim 12, wherein saidfirst sealing assembly and said second sealing assembly each comprise arecessed groove.
 16. A gas turbine engine comprising: a compressor; acombustor coupled downstream from said compressor to receive at leastsome of the air discharged by said compressor; a rotor shaft coupled tosaid compressor; and a plurality of circumferentially-spaced rotorblades coupled to said rotor shaft, each of said plurality of rotorblades comprising: a platform; an airfoil extending radially outwardlyfrom said platform; a dovetail coupled to said rotor shaft; a shankextending between said platform and said dovetail, said shank comprisingat least one cover plate extending inwardly from said platform towardssaid dovetail; and at least one sealing assembly coupled to said coverplate such that a seal path is defined between adjacent rotor blades.17. A gas turbine engine in accordance with claim 16, wherein each ofsaid plurality of rotor blades further comprises a first sealingassembly coupled to said cover plate and an opposite second sealingassembly coupled said cover plate.
 18. A gas turbine engine inaccordance with claim 17, wherein said first sealing assembly comprisesa sealing extension, said second sealing assembly comprises a recessedgroove.
 19. A gas turbine engine in accordance with claim 17, whereinsaid first sealing assembly and said second sealing assembly eachcomprise a sealing extension.
 20. A gas turbine engine in accordancewith claim 17, wherein said first sealing assembly and said secondsealing assembly each comprise a recessed groove.